The present invention relates to a gas turbine engine combustion chamber, more particularly such a combustion chamber having a plurality of axially displaced fuel injection heads.
Gas turbine engines, particularly aircraft turbojet engines, have various modes of operation, in particular a low power mode, a take off mode and a cruising mode. The turbojet engines are required to have low pollution exhaust gases, a requirement which is often in conflict with the requirements to preserve flame stability to avoid combustion chamber flameout during critical flight phases, particularly during landing.
It is known that the production of nitrogen oxides in the exhaust gases are formed at high temperature when the air in the combustion chamber forms a stoichiometric mixture with the fuel and that the production of such nitrogen oxides increases with the amount of time that the stoichiometric mixture dwells in the combustion chamber. These conditions are present in conventional turbojet engines because the combustion chamber volume is quite large in order to assure flame stability at low power operating modes.
In order to reduce the pollutants in the exhaust gases, it has been proposed to provide the combustion chamber with separate fuel injection heads for the low power operating mode and the take off operating mode, which fuel injection heads are radially and axially separated, but supply fuel into a common zone. While this combustion chamber has satisfactorily reduced exhaust pollutants, the radial spacing of the separate fuel injection heads requires the combustion chamber to have an enlarged radial dimension. Furthermore, the large number of fuel injection heads increases both the cost and weight of the turbojet engine.
Another drawback of these known combustion chamber relates to the temperature profile at the combustion chamber outlet when only the lower power fuel injection heads are operating. Since these fuel injection heads are radially spaced from the take off fuel injection heads, the blades of the high pressure turbine, located immediately downstream of the combustion chamber in the exhaust gas flow, are subjected to temperatures near 1,800.degree. K. at their tips and temperature of only approximately 900.degree. K. at their roots. This temperature difference between the blade tips and the blade roots reduces turbine efficiency.